This invention relates generally to gas turbine engines and more particularly to seal bypass circuits in such engines.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. Aircraft engines ordinarily include a stationary turbine nozzle disposed at the outlet of the combustor for channeling combustion gases into the first stage turbine rotor disposed downstream thereof. The turbine nozzle directs the combustion gases in such a manner that the turbine blades can do work.
Because they are exposed to intense heat generated by the combustion process, certain components, such as combustor liners and turbine rotor blades and nozzles, are cooled to meet life expectancy requirements. This cooling is commonly accomplished with relatively cool air that is diverted from the compressor discharge. Typically, a forward outer seal is provided between the stationary turbine nozzle and the first stage turbine rotor for sealing the compressor discharge air that is bled off for cooling purposes from the hot gases in the turbine flow path. Conventional forward outer seals comprise a rotating labyrinth seal made up of a rotating member and a static member that are generally situated circumferentially about the longitudinal centerline axis of the engine. The static member includes an annular flange to which a stator element is mounted. The stator element is normally made of a honeycomb material. The rotating member has a number of thin, tooth-like projections extending radially toward the stator element. The outer circumference of each projection rotates within a small tolerance of the stator element, thereby effecting sealing between the cooling air and the hot gases in the turbine flow path.
During engine operation, certain engine structure does not heat up as fast as other structure because of differences in mass and the degree of exposure to the hot gases. This effect results in radial thermal gradients in many engine components, such as the static member in which the annular flange does not heat up as fast as other portions of the member. Thermal gradients in the static member can cause high thermal stresses and improper seal clearance between the stator element and the rotating tooth-like projections. To reduce the thermal gradient of the flange, it is known to provide a flange bypass circuit through which some of the compressor discharge air bled off for cooling purposes passes. Although this air is cooler than the hot gas flow, it is warm enough to provide faster heating of the flange. The faster heating results in a smaller thermal gradient. Conventional bypass circuits utilize a series of discrete circuits spaced along the circumference of the flange and the stator element's backing plate. This discrete circuit arrangement provides non-uniform heating of the flange, requires elaborate machining, and uses small fillet radii that create stress concentrations.
Accordingly, there is a need for a bypass circuit that would provide more uniform heating of the flange and allow simpler machining.